Technical Field
The present invention relates generally to gas turbine engine turbine blade cooling and, more specifically, cooled turbine blades and slots for mounting the blades.
Background Information
Turbine blades in gas turbine engine turbines and, particularly, high pressure turbine blades are often cooled by a portion of pressurized air from a compressor of the engine. Each turbine stage includes a row of turbine rotor blades extending radially outwardly from a supporting rotor disk with the radially outer tips of the blades being mounted inside a surrounding turbine shroud. Typically, turbine rotor blades of at least the first turbine stage are cooled by the bled portion of the pressurized air from the compressor. The blades include roots slid into and secured by axial slots in a turbine disk.
The blades are typically cooled using a portion of high pressure compressor discharge air bled (also known as compressor discharge pressure or CDP air) from the last stage of the compressor. The air is suitably channeled through internal cooling channels inside the hollow blades and discharged through the blades in various rows of film cooling holes from the leading edge and aft therefrom, and also typically including a row of trailing edge outlet holes or slots on the airfoil pressure side.
Blade cooling air is gathered and transferred from static portions of the engine to the rotating disk supporting the blades. The cooling air passes through the slot and into the blade root from where it is distributed through a cooling circuit having cooling passages in an airfoil of the blade.
The typical turbofan aircraft engine initially operates at a low power, idle mode and then undergoes an increase in power for takeoff and climb operation. Upon reaching cruise at the desired altitude of flight, the engine is operated at lower or intermediate power setting. The engine is also operated at lower power as the aircraft descends from altitude and lands on the runway, following which thrust reverse operation is typically employed with the engine again operated at high power. In the various transient modes of operation of the engine where the power increases or decreases, the turbine blades heat up and cool down respectively.
A slot bottom of the disk is exposed to blade cooling air during engine operation. The cooling air increases the thermal response of the slot bottom creating a large thermal gradient between the slot bottom and bore of the disk. This gradient creates large thermal stresses in both the acceleration and deceleration of the engine. These large thermal stresses reduces the low cycle fatigue life of the disk.
Accordingly, it is desired to provide a gas turbine engine having turbine blade cooling with a design which reduces a thermal gradient in a bottom of a root mounting slot. It is further desired to reduce large thermal stresses in the bottom of the root mounting slot caused by the thermal gradient. It is also desired to increase the low cycle fatigue life of the disk by reducing these thermal stresses.